Andy P. Broeren and Mark G. Potapczuk

Glenn Research Center, Cleveland, Ohio

James T. Riley

FAA William J. Hughes Technical Center, Atlantic City Airport, New Jersey 

Philippe Villedieu

ONERA, The French Aerospace Laboratory, Toulouse, France

Frédéric Moëns

ONERA, The French Aerospace Lab, Meudon, France

Michael B. Bragg

University of Illinois at Urbana-Champaign, Urbana, Illinois 

Swept-Wing Ice Accretion 

Characterization and Aerodynamics

NASA/TM—2013-216555

September 2013

AIAA–2013–2824
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Andy P. Broeren and Mark G. Potapczuk

Glenn Research Center, Cleveland, Ohio

James T. Riley

FAA William J. Hughes Technical Center, Atlantic City Airport, New Jersey 

Philippe Villedieu

ONERA, The French Aerospace Laboratory, Toulouse, France

Frédéric Moëns

ONERA, The French Aerospace Lab, Meudon, France

Michael B. Bragg

University of Illinois at Urbana-Champaign, Urbana, Illinois 

Swept-Wing Ice Accretion 

Characterization and Aerodynamics

NASA/TM—2013-216555

September 2013

AIAA–2013–2824

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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sponsored by the American Institute of Aeronautics and Astronautics

San Diego, California, June 24–27, 2013
Acknowledgments

The authors gratefully acknowledge the support of their respective organizations in this collaborative research effort. Specific 

contributions to this paper were provided by the University of Illinois under NASA Cooperative Agreement NNX12AB04A and 

FAA Cooperative Agreement 10-G-004. The authors thank Jeff Diebold, Gustavo Fujiwara, Brock Wiberg and Brian Woodard 

at University of Illinois and Cris Bosetti, Abdi Khodadoust, Adam Malone, Ben Paul and John Vassberg at the Boeing Company 

for their contributions to this paper. Eric Kreeger and Mary Wadel at NASA Glenn Research Center provided additional technical 

and management review. The NASA-support portion of this research effort is funded under the Atmospheric Environment Safety 

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2013-216555 1

Swept-Wing Ice Accretion 

Characterization and Aerodynamics

Andy P. Broeren and Mark G. Potapczuk

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

James T. Riley

FAA William J. Hughes Technical Center

Atlantic City Airport, New Jersey 08405 

Philippe Villedieu

ONERA, The French Aerospace Lab

F-31055, Toulouse, France

Frédéric Moëns

ONERA, The French Aerospace Lab

F-92190, Meudon, France

Michael B. Bragg

University of Illinois at Urbana-Champaign

Urbana, Illinois 61801

Abstract

NASA, FAA, ONERA, the University of Illinois and Boeing have embarked on a significant, 

collaborative research effort to address the technical challenges associated with icing on large-scale, 

three-dimensional swept wings. The overall goal is to improve the fidelity of experimental and 

computational simulation methods for swept-wing ice accretion formation and resulting aerodynamic 

effect. A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic 

experiments and computational simulations. As the baseline, full-scale, swept-wing-reference geometry, 

this research will utilize the 65 percent scale Common Research Model configuration. Ice-accretion 

testing will be conducted in the NASA Icing Research Tunnel for three hybrid swept-wing models 

representing the 20, 64 and 83 percent semispan stations of the baseline-reference wing. Threedimensional measurement techniques are being developed and validated to document the experimental 

ice-accretion geometries. Artificial ice shapes of varying geometric fidelity will be developed for 

aerodynamic testing over a large Reynolds number range in the ONERA F1 pressurized wind tunnel and 

in a smaller-scale atmospheric wind tunnel. Concurrent research will be conducted to explore and further 

develop the use of computational simulation tools for ice accretion and aerodynamics on swept wings.

The combined results of this research effort will result in an improved understanding of the ice formation 

and aerodynamic effects on swept wings. The purpose of this paper is to describe this research effort in 

more detail and report on the current results and status to date.

Nomenclature

c chord length 

s surface distance measured from leading edge

x chordwise coordinate
NASA/TM—2013-216555 2

y spanwise coordinate (vertical coordinate for airfoils)

z vertical coordinate

AR aspect ratio

Cl sectional lift coefficient

Cd sectional drag coefficient

CP pressure coefficient

M Mach Number

Re Reynolds Number

V velocity

α angle of attack

λ taper ratio

Λ sweep angle

CFD Computational Fluid Dynamics

CRM Common Research Model

CRM65 Common Research Model geometry at 65 percent scale

FAA Federal Aviation Administration

IRT Icing Research Tunnel

LWC Cloud liquid water content

MVD Median volumetric diameter

NACA National Advisory Committee for Aeronautics

NASA National Aeronautics and Space Administration

ONERA Office National d’Etudes et de Recherches Aérospatiales

RPM Rapid Prototyping Manufacturing

WSU Wichita State University

1.0 Introduction

Ice accretion and its aerodynamic effect on highly three-dimensional swept wings are extremely 

complex phenomena important to the design, certification and safe operation of small and large transport 

aircraft. There is increasing demand to balance trade-offs in aircraft efficiency, cost and noise that tend to 

compete directly with allowable performance degradations over an increasing range of icing conditions.

Computational fluid dynamics codes have reached a level of maturity that they are being proposed by 

manufacturers for use in certification of aircraft for flight in icing. However, sufficient high-quality data 

to evaluate their performance on iced swept wings are not currently available in the public domain.

Significant knowledge gaps remain for swept-wing geometries and supercooled, large-droplet icing 

conditions including freezing drizzle and freezing rain. NASA, in collaboration with FAA, ONERA and 

University of Illinois, has developed a multi-phase research effort to address some of these knowledge 

gaps. This paper provides an overview of this research effort and a summary of progress to date.

This research effort is modeled after a previous successful international collaboration that 

investigated aerodynamic effects and ice accretion simulation for airfoils and straight wings (Refs. 1 to 

18) The overall goal of the previous collaboration was to provide high-fidelity, full-scale, iced-airfoil 

aerodynamic data and validated subscale-model simulation methods that produce the essential full-scale 

aerodynamic characteristics. The research was organized into six phases involving icing wind tunnel and 

aerodynamic wind tunnel experiments with both subscale and full-scale models using the NACA 23012 

airfoil. An important initial step in this research was the classification of ice accretion according to the 

aerodynamic effect on the flowfield. This provided a framework within which the ice accretion and 

aerodynamic testing was conducted. The NASA IRT was used to generate the ice accretion using both 
NASA/TM—2013-216555 3

subscale and full-scale models. Aerodynamic testing was performed at the ONERA F1 pressurized wind 

tunnel using a 72 in. (1.83 m) chord, full-span, unswept, NACA 23012 airfoil model with high-fidelity, 

three-dimensional castings of the IRT ice accretions. Performance data were recorded over a large range 

of Reynolds number, up to 15.9×106 and Mach numbers from 0.10 to 0.28. This generated a large, highquality, benchmark, iced-airfoil aerodynamic database. Lower-fidelity simulation methods were 

developed and tested on an 18 in. (0.46 m) chord NACA 23012 airfoil model at the University of Illinois 

wind tunnel at lower Reynolds number. The aerodynamic accuracy of the lower-fidelity, subscale ice 

simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a 

factor of 8 reduction in Reynolds number. Methods were developed for designing lower-fidelity artificial 

ice shapes for subscale models. The completed research defined the level of geometric fidelity required 

for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty.

This previous work has led to a more complete understanding of ice contamination aerodynamic effects 

on airfoils. This is an important building block, but a fundamental question remains as to how relevant 

these 2-D data are to 3-D swept-wing geometries.

Swept-wing icing has been a significant research area of interest for many years. This work has 

mostly focused on icing physics studies aiding in the development of computational simulation tools

(Refs. 19 to 30). Icing physics studies have been conducted to understand the physics of formation of ice 

accretions on swept wings and to develop models that allow their prediction. The development of 

computational simulation tools centers on the continuous improvement of codes such as LEWICE3D 

which predict particle trajectories and ice accretions for 3-D configurations. While this work is critical to 

understanding the relevant physical factors, it has mostly been conducted on small-scale geometries, often 

using NACA 0012 airfoil sections or circular cylinders. In order to continue the advancement of 3-D icing 

simulation tools, new validation data are needed for large-scale, modern, swept-wing configurations.

Relevant 3-D ice accretion geometries must be documented in order to continue the validation efforts for

the current array of icing simulation tools.

There exist in the public domain a small number of aerodynamic studies for iced-swept wings. A 

fundamental study of the flowfield on a subscale rectangular swept wing using a NACA 0012 airfoil 

section was conducted with a simulated, glaze-ice accretion (Refs. 31 to 34) The 3-D velocity 

measurements on the iced wing were complemented with CFD simulations to develop a good 

understanding of the ice-shape effect on the flowfield. In 2001, a research program involving FAA, 

NASA and WSU was conducted to develop an experimental database of ice accretion effects on 

aerodynamic performance of a finite swept wing (Refs. 35 to 37). An icing test was conducted in the 

NASA IRT on a subscale wing having 28° sweep and GLC-305 airfoil section. High-fidelity ice casting 

simulations were generated for aerodynamic testing in the WSU 7- by 10-ft wind tunnel at a Reynolds 

number of 1.8×106

. While this study utilized high-fidelity artificial ice shapes, the small model scale and 

low-Reynolds number limit the applicability of the results to a full-scale airplane. Reehorst et al. (Refs. 38

and 39). also conducted a low-Reynolds number aerodynamic study of icing effects on a realistic, fully 

three-dimensional, 12.5 percent scale model of a twin-engine, short-haul commercial transport. The 

authors considered various roughness size scaling approaches for the small-scale and low-Reynolds 

number investigation. Despite the low-Reynolds number at which the data were acquired, this effort

resulted in a large database of icing effects on a full-aircraft configuration. Aerodynamic testing of a fullscale, swept-wing, business jet, T-tail model was performed under the auspices of the NASA/FAA 

Tailplane Icing Program, Phase II (Ref. 40 and 41). While the T-tail model results are, in general, 

certainly applicable to the swept-wing icing aerodynamics research described here, no 3-D ice-shape 

configurations were tested. 

The very brief literature review contained herein has cited a number of significant research efforts in 

swept-wing ice accretion characterization and aerodynamics. There are a number of low-Reynolds 

number aerodynamic databases for relatively simple geometries. The ice-accretion research conducted has 

also led to a good fundamental understanding of the important physical processes, but mostly at small 

scale. There is very little ice accretion geometry information for large-scale, modern design, swept wings

in the public domain known to the authors. For example, the mean aerodynamic chord length can range 
NASA/TM—2013-216555 4

from approximately 10 ft (3 m) for regional jets to 30 ft (9.1 m) for wide-body airplanes. Similarly, the 

authors are also unaware of any high-Reynolds number aerodynamic performance and flowfield data 

illuminating the effect of ice on swept-wing aerodynamics that is available in the public domain.

Three-dimensional computational and experimental icing simulation tools are becoming increasingly 

common in icing analysis for modern transport airplanes. Manufacturers are proposing the use of CFD 

codes for certification of aircraft for flight in icing conditions. However, validation data are very limited, 

particularly for large-scale swept wings that are typical of modern commercial airplanes. Thus it is 

difficult to determine how much confidence can be placed in results from CFD codes used in design, and 

particularly in certification. Furthermore, an understanding of the icing effects on swept-wing 

aerodynamics is critical to evaluating the accuracy to which ice accretion must be predicted by 

computational tools or simulated in aerodynamic testing. For example, it is known for some cases that 

swept-wing ice accretion can be highly three dimensional (e.g., “scallops” or “lobster tails”), which is 

discussed below in Section 2.1. An important question is how much detail of that three-dimensionality is 

critical to the aerodynamic performance and therefore must be accurately simulated. In addition, basic 

swept-wing aerodynamic effects such as the spanwise flow and configuration dependence play an 

important role. An understanding of scale effects, including Reynolds and Mach number, is needed in 

order to develop lower cost aerodynamic test techniques for iced swept wings. Valid tests conducted at 

smaller scale and lower Reynolds number hold potential for developing a more complete understanding of 

the aerodynamics. Both full-scale and subscale research is needed to develop and validate computational

fluid dynamics (CFD) simulation tools used to predict the aerodynamics of iced-wing configurations. All 

of these factors have provided motivation for collaborative research effort in this arena.

NASA, FAA, ONERA, the University of Illinois and Boeing, have embarked on a significant, 

collaborative research effort to address the technical challenges associated with swept-wing icing. The 

overall goal of ensuring continued flight safety will be achieved by improving the fidelity of experimental 

and computational simulation methods for swept-wing ice accretion formation and resulting aerodynamic 

effect. There are three specific objectives:

• Generate a database of 3-D, swept-wing, ice-accretion geometries for icing-code development 

and validation and for aerodynamic testing.

• Develop a systematic understanding of the aerodynamic effect of icing on swept wings including: 

Reynolds and Mach number effects, important flowfield physics and fundamental differences 

from 2-D.

• Determine the level of ice-shape geometric fidelity required for accurate aerodynamic simulation 

of swept-wing icing effects.

A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic 

experiments and computational simulation to address these objectives. This is depicted schematically in 

Figure 1. Phase I is a review of the technical literature associated with iced swept-wing aerodynamics 

with emphasis on classifying ice accretion based upon key aerodynamic features of the flowfield. This 

research provides an organizational framework for the icing and aerodynamic experiments conducted in 

later phases. The research conducted in Phase II will identify the baseline swept-wing model to be used 

for the experiments and computational analysis. In addition, measurement techniques will be developed to 

document, in 3-D, the experimental ice-accretion geometries. Aerodynamic measurement techniques

suitable to iced swept-wing experiments will also be investigated. The icing tests conducted in Phase III 

will generate a large database of ice-accretion geometries on realistic large-scale swept wings. Artificial 

ice shapes of varying geometric fidelity will be developed for aerodynamic testing over a large Reynolds 

number range in Phases IV and V. Exploring the effects of geometric fidelity and Reynolds number will 

require final validation tests in Phase VI. The research conducted in Phase VII will explore and further 

develop the use of computational simulation tools for ice accretion and aerodynamics on swept wings.
NASA/TM—2013-216555 5

Figure 1.—Seven-phase research effort designed to investigate large-scale swept wing ice 

accretion and aerodynamics.

Each phase has a number of technical challenges that must be overcome in order for the research 

objectives to be achieved. The purpose of this paper is to describe these technical challenges in more 

detail and report on the current results and status to date.

2.0 Initial Research Results and Plans

2.1 Phase I: 3-D Ice Accretion Classification

Ice accretion is often described in terms of its appearance or composition, such as glaze, rime, clear, 

mixed, runback, beak, and intercycle among others. While this terminology may be useful for 

characterizing the visual attributes of ice accretion, it may not be as useful when the objective is to 

understand the aerodynamic effects. Bragg et al. (Refs. 3 and 4) examined the icing aerodynamics 

literature and developed four fundamental types or categories based upon the flowfield physics that was 

unique to each category. This research was directed at the essentially 2-D aspects of iced airfoils.

Categorizing ice accretion in this way provided an aerodynamic framework for the research designed to 

quantify the effects of geometric fidelity on iced-airfoil aerodynamics. This proved to be a successful 

approach and was identified as an important part of the swept-wing icing research effort.

Ice accretion formations on swept wings can have unique characteristics. Depending upon specific 

icing conditions and sweep angle, the region of the attachment line on the ice accretion may not be 

smooth as is often the case for straight wings. This is depicted in Figure 2 that shows an example of initial 

roughness and a large rime-ice accretion on a swept wing. For glaze icing, certain combinations of icing 

conditions and sweep angle can lead to the formation of highly 3-D features called “scallops” or “lobster 

tails” that do not exist for ice accretions on airfoils. An example of this type of swept-wing ice accretion 

is shown in Figure 3. It is also possible to have glaze ice accretion with no scallops or incomplete scallop 

formations on swept wings (Ref. 19).
NASA/TM—2013-216555 6

Figure 2.—Photographs of initial roughness (left) and rime ice (right) accreted on a NACA 0012 

wing having 45° leading-edge sweep in the NASA Glenn Icing Research Tunnel.

Aerodynamic and icing conditions were V = 200 kn, α = 0°, LWC = 0.45 g/m3

, MVD = 32 μm, 

total temperature = 20 °F (left) and 1 °F (right), exposure time = 2 min (left) and 20 min (right); 

adapted from Broeren et al. (Ref. 42).

Figure 3.—Photographs of complete scallop glaze ice accreted on a NACA 0012 wing having 

45° leading-edge sweep in the NASA Glenn Icing Research Tunnel. Aerodynamic and icing 

conditions were V = 200 kn, α = 0°, LWC = 0.45 g/m3

, MVD = 32 μm, total temperature = 

20 °F, exposure time = 20 min; after Broeren et al. (Ref. 42).
NASA/TM—2013-216555 7

Broeren et al. (Ref. 42) have prepared an extensive review of the existing data on swept-wing ice 

accretion and aerodynamics. The existing data tend to be: (1) mostly at low-Reynolds number and (2) 

applicable to simple swept-wing geometries that do not have high-lift systems, wing-mounted engines, 

fuselages and other features of actual airplane wings. These factors can significantly alter the iced 

aerodynamics for particular configurations and so extreme caution must be exercised in terms of making 

general conclusions based upon the current, limited database. Given what is known from the existing data, 

the same four classifications, or fundamental categories used for iced airfoils were suggested: (1) 

roughness, (2) horn ice, (3) streamwise ice and (4) spanwise-ridge ice. Instead of relying upon iceaccretion terminology such as rime and glaze, the four aerodynamic groups have names associated with 

ice-shape geometry. Broeren et al. (Ref. 42) describe the unique flowfield features of each group that 

determine the iced-swept-wing aerodynamics.

• Ice roughness represents initial leading-edge ice accretion and a key aerodynamic characteristic is 

that the scale of the boundary-layer separation is of the same order as the size of the roughness.

How this small, separated flow region interacts with the spanwise flow is unknown.

• Horn ice is large, leading-edge ice accretion that can be associated with glaze icing conditions.

The flowfield is characterized by large-scale, boundary-layer separation originating at the horn.

This separation leads to the formation of a spanwise-running, leading-edge vortex that is similar 

to that found on clean swept wings with leading-edge separations. The small amount of existing 

data indicate that there are fundamental flowfield differences between nominally 3-D horn ice 

characterized by no scallop formations versus highly 3-D horn ice characterized by fully 

developed scallop formations. Diebold et al. (Ref. 43) provide a more detailed description of this 

complex flowfield.

• Streamwise ice can be associated with rime icing conditions and is generally conformal to the 

wing leading edge, or may form a horn-like feature (or protuberance) oriented into the flow 

direction. The only example of this group cited by Broeren et al. (Ref. 42) showed an increase in 

wing maximum lift coefficient with the streamwise ice. While this effect is thus known to be 

possible, it is not expected to hold for most cases and illustrates the need for further wing 

performance data and flowfield information with realistic streamwise ice simulations.

• Spanwise-ridge ice can be associated with droplet impingement and ice formation aft of the area 

covered by an ice-protection system in supercooled, large-droplet icing conditions or with

incomplete evaporation of impinging water. The leading edge is free of ice with an ice ridge 

located downstream often in the range of 10 to 15 percent chord. Broeren et al. (Ref. 42) describe 

data from only one low-Reynolds number study for very simple geometric representations of 

spanwise-ridge ice on a swept wing.

The lack of data noted for streamwise and spanwise-ridge ice was also true to a lesser degree for 

roughness and horn ice. The result was an inability to fully develop the classification system. Therefore, 

Broeren et al. (Ref. 42) make a number of specific recommendations where more data are needed and 

thus provide additional guidance for this research effort. It is likely that the classification system will be 

revisited as more data become available. In this sense, the Phase I research is an ongoing effort 

throughout the course of the project.

2.2 Phase II: Ice Accretion and Aerodynamic Measurement Methods Development

The research conducted in Phase II provides the necessary foundation for the experimental and 

computational work to be conducted in the latter phases. The Phase II research is organized into three 

areas: 

1. Defining the baseline, full-scale, swept-wing model geometry to be used for the research effort;

2. Developing and validating methods to measure highly 3-D ice accretion;
NASA/TM—2013-216555 8

3. Developing the appropriate measurement methods to quantify the iced-swept wing aerodynamics.

Significant progress has been achieved in each of these areas and is briefly described here.

2.2.1 Baseline Model Selection

For this research to be useful and relevant, it is important to select a baseline, swept-wing model 

geometry that is representative of current, modern design civilian transport airplanes. The selection 

process was complex given the large number of variables, such as sweep angle, aspect ratio, mean 

aerodynamic chord and wing span. An additional requirement was for all of the geometry to be nonproprietary and non-export controlled. After reviewing the available options, the Common Research 

Model geometry was selected (Ref. 44). The Boeing Company provided the design of the CRM for a

previous joint experimental effort with NASA which fabricated test articles and conducted aerodynamic 

testing (Refs. 45 and 46). The CRM geometry was also used as part of the AIAA Drag Prediction 

Workshops (Ref. 47). The model, shown in Figure 4 with and without the engine nacelle/pylon, includes a 

fuselage that is representative of a wide-body commercial transport airplane. The 35° swept wing has a 

contemporary transonic supercritical design that is well behaved with and without the nacelle/pylon 

structure. Table I provides a comparison of the CRM wing geometry with existing wide-body airplanes. A 

key advantage to the CRM is that all of the geometry information, CFD analysis and experimental data 

are available in the public domain. This fact, coupled with its modern design, make the CRM an ideal 

baseline reference model for this research.

The fact that the CRM is representative of a wide-body transport airplane means that the physical size 

is very large compared to that of many other swept-wing airplanes such as single-aisle commercial 

transports including regional and business jets. The large physical size of the CRM wing presents specific 

challenges to both the ice-accretion and aerodynamic testing. Large-scale wing ice-accretion testing 

requires the design of “hybrid” or “truncated” models where the full-scale leading-edge geometry is 

matched to a shortened or truncated afterbody. This design process is the subject of Phase III of the 

Figure 4.—Conceptual design of the Common Research Model with and without engine nacelle/pylon, after Vassberg 

et al. (Ref. 44).

TABLE I.—COMPARISON OF CRM WING GEOMETRY WITH EXISTING WIDE-BODY AIRPLANES(a)

Airplane Span,

ft

Mean aerodynamic 

chord,

ft

Area,

ft2

Aspect

ratio(b)

Taper

ratio(b)

Sweep

angle, c/4

CRM 192.8 23.0 4,130 9.0 0.28 35°

Airbus A330-200/300 198.0 23.9 3,892 9.5 0.22 30°

Boeing 777-200 199.9 26.5 4,389 8.7 0.27 31°

Boeing 787-9 197.0 20.6 3,880 9.6 0.18 32°

Boeing 747-400 211.4 29.8 5,417 7.7 0.28 37° (a) Data for existing wide-body airplanes was compiled from publically available sources that may use different conventions to define the 

geometric parameters. 
NASA/TM—2013-216555 9

research effort described in Section 2.3. The large physical size of the CRM wing requires a very 

aggressive design for the hybrid model that presents greater risk of adverse effects when installed in the 

NASA IRT. For aerodynamic wind-tunnel testing, the full-scale CRM wing can be scaled to an 

appropriate size for the facility. In the case of the ONERA F1 wind-tunnel, an 8 percent scale model of 

the CRM would be of appropriate size. While this is not unreasonable for the clean, baseline model 

geometry, such a large scale reduction becomes challenging when the goal is to accurately simulate iceaccretion geometry that typically includes small roughness. Typical ice-roughness sizes on the full-scale 

model ice accretion could be in the range of 0.04 to 0.08 in. (1 to 2 mm) which is equivalent to 0.003 to 

0.006 in. (0.08 to 0.16 mm) on the 8 percent scale aerodynamic model. Small roughness features of this 

size are very challenging to accurately reproduce on the artificial ice shapes developed for aerodynamic 

testing. Based upon these factors, the research team decided to use a 65 percent scale version of the CRM 

as the full-scale, baseline, reference geometry for this research. Table II provides a comparison of the 

CRM65 geometry with that of existing single-aisle commercial transport airplanes. The geometries are 

similar in scale with only the CRM65 sweep angle being considerably larger. More detailed geometry 

information for the CRM65 wing semispan is shown in Figure 5. Using the CRM65 as the full-scale, 

baseline, reference geometry for this research reduces potential risks associated with the ice-accretion and 

aerodynamic testing while still being representative of current transport airplanes.

TABLE II.—COMPARISON OF CRM65 WING GEOMETRY WITH EXISTING SINGLE-AISLE AIRPLANES(a)

Airplane Span,

ft

Mean aerodynamic 

chord,

ft

Area,

ft2

Aspect 

ratio(b)

Taper 

ratio(b)

Sweep 

angle, c/4

CRM65 125.3 15.0 1,745 9.0 0.28 35°

Airbus A320 112.0 14.1 1,320 9.5 0.21 25°

Boeing 737-800 112.6 13.0 1,341 9.5 0.16 25°

Boeing 757-200 124.8 16.7 1,847 7.8 0.21 25° (a) Data for existing single-aisle airplanes was compiled from publically available sources that may use different conventions to define the 

geometric parameters.

Figure 5.—Summary of CRM65 wing geometric characteristics, adapted from Vassberg et al. (Ref. 44)

Yehudi break, 37% semispan

Fuselage side of body, 10% semispan

Symmetry plane

CRM 65 Wing Geometry

• Semispan = 62.7 ft

• Root chord (symmetry plane) = 29.0 ft

• Root chord (fuselage side of body) = 25.4 ft

• Tip chord = 5.8 ft

• Mean aerodynamic chord = 15.0 ft

• Semispan area = 873 ft2

• Aspect Ratio = 9.0

• Taper Ratio = 0.28

• Sweep angle (c/4) = 35°
NASA/TM—2013-216555 10

2.2.2 3-D Ice Accretion Measurement

Generating a database of 3-D experimental ice-accretion geometries and associated artificial ice 

shapes for subsequent aerodynamic testing requires a robust measurement system. Standard ice-accretion 

documentation methods are generally 2-D, such as cross-sectional tracings and qualitative, such as 

photographs. While these methods have been used for many years, they have significant shortcomings 

when used for the potentially highly 3-D ice accretions that can occur on swept wings (cf. Figure 2 and 

Figure 3). For example, Bosetti et al. (Ref. 48) show a comparison of three dissimilar tracings of the same 

ice shape along the span of a swept wing. The best current technology for capturing 3-D features of ice 

accretion is the mold and casting method. This has been used for many years using various materials and 

was improved at NASA Glenn during the 1980s using more robust materials to improve accuracy and 

durability (Ref. 49). A significant disadvantage to this method is that there is no digitized record of the ice 

accretion. Therefore, the present objective is to develop a method to accurately and efficiently digitize ice 

accretion in 3-D. Furthermore, the data must be processed and archived so that: (1) comparisons to iceaccretion code results can be performed; (2) artificial ice shapes can be readily fabricated for aerodynamic 

testing; and (3) the geometry may be readily adapted for CFD simulations.

NASA is leading the effort to adapt commercial laser-based scanning methods that can quantify the 

full three-dimensional features of ice accretion in the IRT. Lee et al. (Ref. 50) describe the significant 

progress that has already been made in accomplishing this task. Demonstration tests of several different 

scanning systems and software were conducted in the IRT and evaluated against a set of pre-defined 

criteria. It was found that the scanning technology and capability was similar over the range of systems 

evaluated. An articulated-arm based system was selected because it required minimal modifications to the

existing IRT test-section resulting in fewer risks during use.

Several different software packages were considered for post-processing of the data. The processing 

steps included combining individual scans, developing a surface mesh and filling holes or gaps in the scan 

data. The most critical software function for working with ice scan data is the ability to create closed, 

watertight surfaces of the highly irregular, rough features of ice accretion. Through a series of hardware 

demonstrations, it was determined that the Geomagic software package has the capability to process 

irregular “organic” surfaces like those typical of ice accretion. The software is also able to create data 

files that can be used for various rapid prototyping manufacturing methods. Such methods can be used to 

fabricate artificial ice shapes from the scan data. The software also has exact surfacing capability that can 

be used to develop grids for computational analysis. Having selected and purchased the scanner and 

software system, the next step was conducting validation research to define the capability.

The scanner system validation research was divided into two parts: one for straight wing ice shapes

and another part dedicated to swept-wing ice shapes. Both of these activities are currently underway. In 

the former case, an aerodynamic validation will be conducted. Ice-accretion testing has been completed 

using an 18-in. chord NACA 23012-airfoil, straight-wing model in the IRT. Several different ice 

accretions were generated ranging from small roughness to large, glaze-horn shapes. The ice accretion 

was measured with the scanning system and then a mold was made of that ice accretion. Subsequently, 

artificial ice shapes were fabricated using the casting method from the molds and from RPM methods 

based upon the 3-D scan data. These artificial ice shapes will be mounted to the leading edge of a NACA 

23012 airfoil model for aerodynamic testing at the University of Illinois Low-Speed, Low-Turbulence 

wind tunnel. The result will be an aerodynamic effect comparison between artificial ice shapes produced

from molds and castings vs. shapes produced from 3-D scans and RPM methods. Figure 6 shows a 

comparison of pencil tracing of the ice castings to cross-sections extracted from the scan data. A common 

feature in the comparison is that the pencil tracings tend to be much smoother than the scan data. This was 

expected since the pencil tracing does not have high enough resolution to capture small geometry 

variations in the ice accretion. More significant differences such as that observed for the ED1978 case on 

the upper surface near x/c = 0.0 can usually be attributed to spanwise variation in the ice accretion 

geometry combined with small differences in the spanwise location of a pencil tracing vs. the scan-data 

cross section. Overall, there is excellent agreement in Figure 6 between the digitized pencil tracings and 

the cross-sections extracted from the scan data.
NASA/TM—2013-216555 11

Figure 6.—Comparison of ice accretion cross-sections taken from digitized pencil tracings and 

3-D digital scan data for four IRT runs (ED1966, ED1967, ED1977 and ED1978) on an 18-in.

chord NACA 23012 straigt wing model.

The validation plan for swept-wing ice shapes includes a geometric comparison between the 3-D laser 

scan data and 3-D obtained from other methods such as commercial computed tomography (CT)

scanning. A significant challenge associated with swept-wing ice accretion is obtaining scan data within 

the highly 3-D features such as “scallops” or “lobster tails.” The arm-based, laser-scanning system uses a 

“line-of-sight” method that cannot acquire data within small surface gaps. In this case, more manual 

intervention during post-processing is required in order to create closed or watertight surfaces. As shown 

in Figure 7, reasonable qualitative results can be obtained. Non-line-of-sight based scanning methods will 

be investigated to evaluate the accuracy of the laser-based system for highly 3-D ice geometries. The 

results of this work will be presented in future reports and papers.

x/c

y/c

-0.04 -0.02 0.00 0.02 0.04 0.06 0.08 -0.04

-0.02

0.00

0.02

0.04

0.06

NACA 23012

ED1978 Scan

ED1978 Tracing

x/c

y/c

0.00 0.05 0.10 0.15 -0.05

0.00

0.05

0.10

NACA 23012

ED1977 Scan

ED1977 Tracing

x/c

y/c

0.00 0.05 0.10 0.15 -0.05

0.00

0.05

0.10

NACA 23012

ED1966 Scan

ED1966 Tracing

x/c

y/c

0.00 0.05 0.10 0.15 0.20 0.25 -0.10

-0.05

0.00

0.05

0.10

NACA 23012

ED1967 Scan

ED1967 Tracing
NASA/TM—2013-216555 12

Figure 7.—Comparison of 3-D scan data (left) with photograph (right) of an IRT ice accretion, 

after Lee et al. (Ref. 50).

2.2.3 Aerodynamic Measurement Methods

An aerodynamic framework for classifying swept-wing ice accretion based upon unique flowfield 

features was described in Section 2.1. Identifying these flowfield features and their contribution to the 

resulting aerodynamic performance degradation is key to satisfying the objectives of this research effort

as described in the Introduction (Section 1.0). In addition to analyzing standard performance data such as 

lift, drag and pitching moment, it is important to understand the characteristics of the flowfield that drive

the changes in performance. Flowfield information is also required for the continued development and 

validation of computational simulation tools. The development of flowfield measurement methods 

contributes to the planned high- and low-Reynolds number testing of the scaled CRM swept wing with 

artificial ice shapes.

This methods development effort has been carried out at small scale and very low-Reynolds numbers 

at the University of Illinois. Diebold et al. (Refs. 51 to 53) describe the application of pressure sensitive 

paint and wake survey techniques to iced swept wings. These experiments were conducted on an 

approximately 2 percent-scale, modified CRM wing model with an artificial leading-edge ice shape.

Figure 8 shows a comparison of pressure-sensitive paint results with surface-oil flow visualization on the 

iced wing upper surface. The iced wing is stalled at this angle of attack. The flow visualization indicates 

that there was a spanwise-running, leading-edge vortex on the inboard portion of the wing. The 

reattachment line (also shown in Figure 9) is evident in the sharp streamwise pressure gradient between 

the low- and high pressure regions (centered near Cp ≈ –1.2) measured using pressure-sensitive paint. The 

surface-pressure data have the ability to quantify the three-dimensional flow separation features observed 

in the oil-flow patterns. These results are especially encouraging given the known challenges associated 

with pressure-sensitive paint measurements at low-dynamic pressure.

The wake flowfield surveys were conducted using a 5-hole pressure probe to yield total and static 

pressure in addition to all three velocity components. Shown in Figure 9 is again the surface-oil flow 

visualization image from Figure 8 along with the velocity measured in the wake. In this figure, the 

regions of significant flow separation on the iced wing are correlated to regions of higher streamwise 

momentum deficit in the wake. For example, consider the stalled region on the wing nominally located at 

0.6 < 2y/b < 0.8 which the wake survey data have quantified. It is expected that the contribution to the 

iced wing profile drag measured in this region would be significant.
NASA/TM—2013-216555 13

Figure 8.—Comparison of surface oil flow visualization (left) with 

pressure-sensitive paint (right) results for scale swept wing with 

artificial ice shape at α = 6.5°, Re = 0.6×106 and M = 0.15, after 

Diebold et al. (Ref. 51).

Figure 9.—Comparison of surface oil flow visualization (top image) with 

3-D wake surface data; contours of streamwise velocity with vectors 

of transverse velocity for scale swept wing with artificial ice shape at 

α = 6.5°, Re = 0.6×106 and M = 0.15, after Diebold et al. (Ref. 51).

-3 -2.5 -2 -1.5 -1 -0.5 0

CP

2y/b

2z/b

0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 -0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.99

0.97

0.95

0.93

0.91

0.89

u/U∞

Vortex

Attached 

Inboard 

Section 

Secondary 

Separation

Reattachment 

Lines

Stalled

Region
NASA/TM—2013-216555 14

In fact Diebold et al. (Ref. 52 and 53) have demonstrated this to be the case through careful 

integration of the wake survey data. As shown in Figure 10, the integration yields the spanwise 

distributions of lift and both profile and induced drag for the stalled flowfield of the clean and iced wing.

Comparison of the absolute values for the clean and iced configurations is complicated by the significant 

difference in angle of attack. The results for the iced configuration, however, can be directly compared to 

the results in Figure 8 and Figure 9. The spanwise distribution of lift shows a significant decrease in the 

region nominally located at 0.6 < 2y/b < 0.8 identified in Figure 9 as stalled. As indicated, this 

corresponds to a significant increase in the profile drag. Data similar to that shown in Figure 8 to Figure 

10, acquired for realistic ice shapes at reasonable Reynolds number, have the potential to provide the 

flowfield information necessary to understand changes in swept-wing performance due to the artificial ice 

shapes as well as the effects of ice-shape geometric fidelity.

Figure 10.—Comparison of spanwise distributions of lift and drag as measured on the scale swept wing at 

α = 6.5°, Re = 0.6×106 and M = 0.15, after Diebold (Ref. 52).
NASA/TM—2013-216555 15

These methods have been investigated by many researchers and are well established. What is unique 

in this case is the potential challenges presented by the complexities of an iced swept-wing flowfield. For 

airfoils, it was found that the addition of artificial ice geometries often resulted in large-scale unsteady, 

three-dimensional and separated flow. These situations can challenge the implementation of and 

interpretation of results from the selected measurement methods. The research conducted in Phase II will 

work to address these concerns while building a foundation of experience in processing and analyzing 

these data.

2.3 Phase III: Ice-Accretion Testing

Ice-accretion testing will be conducted during Phase III generating an ice-accretion geometry

database to be used for icing code validation and aerodynamic experiments. The design of ice-accretion 

experiments for the IRT is complicated by the typical large scale of swept wings on transport airplanes 

such as the CRM. As previously described in Section 2.2.1, the large chord length of the CRM65 wing 

requires a hybrid design approach in order to generate representative, full-scale ice accretion in the 6- by 

9-ft IRT test section. A hybrid swept-wing model preserves the full-scale leading-edge geometry in order 

to obtain ice accretions representative of the full-scale airplane, but requires a custom-designed aft section 

to maintain the droplet impingement and leading-edge flow characteristics and resulting ice shape. Since 

the span of CRM65 wing is much larger than the IRT test section, certain spanwise sections of the wing 

must be selected for the hybrid model design and subsequent ice accretion testing. The University of 

Illinois-Boeing team is leading the research effort to address these technical challenges to help ensure the 

successful completion of the ice-accretion testing. This work is being conducted under four main tasks:

1. Selection of airplane mission and corresponding flight and icing conditions;

2. Simulation of the full-scale icing conditions;

3. Design and simulation of hybrid model wing sections in 2-D;

4. Design and simulation of hybrid model wing sections in 3-D.

As described in Section 2.2.1, the full-scale flight baseline airplane selected for this work is the

CRM65. The approach in the first two tasks is to generate flowfield and ice-shape information in realistic 

flight and icing conditions. These results, such as the location of the attachment line along the wing, 

collection efficiency and ice-shape profiles then become the reference standard for the 2-D and 3-D 

hybrid model design studies carried out in the latter two tasks. Significant progress has been achieved in 

these areas and is briefly described here. The development of these hybrid model design methods is also a 

significant outcome of this phase of the research effort.

2.3.1 Selection of Airplane Mission and Corresponding Flight and Icing Conditions

A set of icing mission scenarios was defined that were typical of large commercial transport airplanes 

and included climb, cruise, hold and descent phases of flight. The selection of airplane weights, flight 

speeds, altitudes and angles of attack for each flight phase was appropriate for an airplane of the CRM65

class. The selected icing conditions were based upon the Code of Federal Regulations Part 25 Appendix C 

continuous maximum envelope and thus defined droplet MVD, cloud LWC and temperature.

2.3.2 Simulation of Full-Scale Icing Conditions

The large matrix of flight and icing conditions defined in task 1 were reviewed and a smaller number 

of cases were selected for further analysis. This subset of icing scenarios provided for a range of ice 

accretion on the full-scale airplane, while significantly reducing the workload associated with analyzing 

all cases. Flow simulations were performed at each of the selected flight conditions using the 3-D RANS 

code OVERFLOW (Ref. 54), thus generating a large database of flowfield information for the clean, 

flight baseline of the full-scale airplane. The flowfield solution was used as input to the LEWICE3D iceaccretion prediction code to generate ice-shape results for the corresponding flight conditions. Some 
NASA/TM—2013-216555 16

results are shown in Figure 11 and Figure 12 for a flight condition at 10,000 ft altitude, speed = 232 kn, 

static temperature = –4 °C, airplane angle of attack = 3.7°, droplet MVD = 20 μm, cloud LWC = 0.55 

g/m3 and an exposure time of 45 min. The contours of local collection efficiency in Figure 11 show the 

location of water impingement on the nose section of the fuselage and wing leading edge. The close-up 

view near the wing tip illustrates the regions of highest local collection efficiency. Figure 12 depicts the 

LEWICE3D generated ice shapes at several locations along the span of the wing. Close up views are 

shown near the wing root and tip sections. For this case, the predicted ice shape is a large upper-surface 

horn. Results of this type were generated for all of the icing cases selected for analysis.

Figure 11.—LEWICE3D local collection efficiency results for CRM65, droplet MVD = 20 μm, V = 232 kn, 

altitude = 10,000 ft.

Figure 12.—LEWICE3D ice shape results for CRM65, droplet MVD = 20 μm, V = 232 kn, altitude = 

10,000 ft, LWC = 0.55 g/m3

, static temperature = -4 °C, 45-min exposure.
NASA/TM—2013-216555 17

Figure 13.—Normalized horn thickness and horn angle taken from the 

LEWICE3D results shown in Figure 12.

The LEWICE3D results were analyzed to understand the spanwise variation in the predicted ice 

geometry. The maximum thickness of the ice and the angle associated with that maximum thickness 

location were calculated along the span of the wing. For the ice shapes shown in Figure 12, this represents 

the length of the ice horn and the angle of the horn with respect to the chord line at each wing station.

These data are plotted in Figure 13 where the horn length and ice thickness are normalized by the local 

chord length. In this figure, the horn length is measured from the tip of the horn to the center of the 

leading-edge radius whereas the ice thickness is the portion of this length from the wing surface to the 

horn tip. The data show that for this case, the normalized ice thickness tends to increase from the root to 

about 50 percent semispan, where there is a decrease. The normalized ice thickness then tends to increase 

from about 55 percent semispan outward to the wing tip. These plots were generated and scrutinized for 

each set of LEWICE3D cases.

A major outcome of this task was the selection of the spanwise locations to be used for the hybrid 

model designs. Determining the number of spanwise locations to select required a balance of competing 

factors. In this research effort, each of the selected locations will be used as the basis for the design of a 

hybrid swept-wing model to be constructed and tested in the IRT. Selecting a large number of spanwise 

locations provides the highest fidelity representation of the ice accretion along the entire wing. Since the 

hybrid model design process will result in large and complex models, practical considerations of time and 

resources limit the total number that can be successfully constructed and tested in the IRT. Estimates of 

model size, time to complete the hybrid-model designs, fabrication cost, number of IRT test days 

required, and cost of IRT testing were developed. After reviewing these estimates it was decided that 

selecting three spanwise stations for further analysis provided a reasonable balance of the competing 

factors.

An additional factor for selecting three spanwise stations was driven by the expectation that these 

stations should be located near the wing: (1) root, (2) midspan and (3) tip. The section near the wing root, 

or inboard station, was selected to be at 20 percent semispan because this corresponded to the minimum 

horn-ice angle for nearly all of the icing cases analyzed. This is the case for the data shown in Figure 13.
NASA/TM—2013-216555 18

Figure 14.—Sketches of CRM65 wing showing selected hybrid model design stations at 20, 64, 83 percent semispan.

The midspan station was selected to be at 64 percent semispan because this corresponded to the maximum 

horn-ice angle for nearly all of the icing cases analyzed. This is also illustrated in Figure 13. The section 

near the tip, or outboard station, was selected to be at 83 percent semispan. This location approximates the 

outboard extent of the wing leading-edge ice protection system in some cases, thus making the icing 

characteristics significant for that reason. This location is also about halfway between the 64 percent

semispan station and the wing tip. These three wing stations are shown graphically in Figure 14.

2.3.3 Design and Simulation of Hybrid Model Wing Sections in 2-D

The design of hybrid models for icing tests where the full-scale leading-edge geometry is combined 

with a truncated afterbody has been explored by Saeed et al. (Refs. 55 to 57 Past research was conducted

primarily for 2-D wing sections and for moderate model sizes that did not require special consideration of 

tunnel wall interference effects. In the current effort, these 2-D design methods were adapted to the 

swept-wing geometry and extended to include the effects of the tunnel walls on the large-blockage-model 

flowfield. Fujiwara et al. (Ref. 58) provide a detailed description of this process along with the results to 

date. Therefore, only a brief example of interim results is provided here. As the design process is 

currently on-going, these results are likely to change prior to the completion of this task.

In simple infinite-swept wing theory, the flow at a spanwise wing station is approximated by the flow 

about the leading-edge-normal airfoil section at the appropriate 2-D Mach number and angle of attack.

Since 2-D hybrid airfoil design methods are relatively well developed, the first step in the swept-wing 

wind-tunnel model design process takes advantage of these 2-D tools to produce a representative 2-D 

hybrid airfoil. Then, in a later process (Section 2.3.4), this is extended into a 3-D swept-wing model and 

validated using 3-D simulation tools and results.

As an example, the hybrid model design process was conducted for the wing station at 64 percent

semispan. The airfoil used for the analysis was taken perpendicular to the leading edge. A comparison of 

the airfoil geometries and surface pressure coefficient results are shown in Figure 15. The full-scale 

leading edge was maintained in the hybrid airfoil to x/c = 0.05 on the upper surface and x/c = 0.10 on the 

lower surface. These limits were set based upon the expected ice accretion limits from the LEWICE3D 

results on the full-scale airplane described in the previous section. Downstream of these locations the 

hybrid airfoil contour was significantly different from the full-scale airfoil and includes a single-element,

trailing-edge flap. The total chord length of the hybrid airfoil (including the flap) is about half the size of 

the full-scale chord length. The flap plays an important role in the hybrid design method since it can be 

used to adjust the airfoil circulation and thus the leading-edge flowfield (stagnation point, minimum 

pressure coefficient, etc.) and as a result, provides additional means of matching the full-scale conditions.
NASA/TM—2013-216555 19

The pressure coefficient comparison in Figure 15 shows what is considered to be typical matching of the 

hybrid and full-scale surface pressure in the region of the leading edge and stagnation point which is of 

importance for matching the local collection efficiency and resulting ice shape. It is not expected that the 

hybrid and full-scale pressures should match downstream of the leading-edge region since the airfoils are 

significantly different. The goal of the hybrid airfoil design methodology is to provide the best matching 

to the full-scale local collection efficiency and leading-edge flowfield and, as a result, the ice shape. This

comparison is shown in Figure 16 where the agreement is very good for this optimized hybrid design.

These hybrid and full-scale airfoil results were generated using LEWICE (a 2-D code). However, the 3-D, 

full-scale, OVERFLOW and LEWICE3D ice-shape results are also used in the hybrid design process.

Figure 16 shows a comparison of the hybrid model and full-scale, LEWICE3D ice shapes. This example 

illustrates that the hybrid model design process can effectively simulate the full-scale icing results with a 

significantly smaller model. Additional research is underway to address the effects of the tunnel walls in 

both 2-D and 3-D simulations.

Figure 15.—Comparison of full-scale and hybrid airfoil geometry and surface 

pressure distribution for 64 percent semispan station. (Note: these are 

interim results intended as an illustrative example.)

Figure 16.—Comparison of full-scale and hybrid airfoil local collection efficiency and ice shape results for 64 percent

semispan station. (Note: these are interim results intended as an illustrative example.)

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9

-3

-2

-1

0

1

Cp

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

-0.1

0

0.1

0.2

x/c

y/c

Full Scale - α=4°

Hybrid+Flap 11° - α=4°

-0.1 -0.05 0 0.05 0.1 0.15 0.2

0.1

0.2

0.3

0.4

0.5

β

s/c

Full Scale - α=4°

Hybrid+Flap 11° - α=4°

0 0.02 0.04 0.06

-0.03

-0.02

-0.01

0

0.01

0.02

0.03

0.04

y/c

x/c

Clean

Full Scale Iced - α=4°

Full Scale Iced (LEWICE3D) - α=4°

Hybrid+Flap 11° Iced - α=4°
NASA/TM—2013-216555 20

Figure 17.—Conceptual sketch of 3-D hybrid swept-wing model in IRT test section.

2.3.4 Design and Simulation of Hybrid Model Wing Sections in 3-D

Having demonstrated computationally in task 3 that the 2-D hybrid airfoil, coupled with infinite 

swept-wing theory, can accurately generate the ice accretion at a section of a finite swept wing, the fullspan, 3-D IRT model can be designed. In the IRT testing, the 3-D hybrid model is used to generate the ice 

accretion found at one spanwise station of the full-scale swept wing, which significantly simplifies the 

hybrid model design process. The 2-D hybrid airfoil sections are extended and swept to create the 3-D 

models for testing in the IRT. Consistent with infinite-swept-wing theory, these models have zero twist 

and zero taper greatly simplifying the design and ultimately, the construction of the models.

Once the models are designed, they must be validated using sophisticated 3-D computational tools

capable of modeling the aerodynamics, droplet impingement and ice accretion in the realistic, highly 3-D, 

IRT environment. The analysis to be carried out includes 3-D flow simulations along with LEWICE3D 

simulations for the hybrid model in the IRT test section. These results will be evaluated against the fullscale flight baseline results described in Section 2.3.2 to determine if any design changes are needed. A 

conceptual sketch of a 3-D hybrid model in the IRT test section is shown in Figure 17. The large size of 

the model relative to the test-section dimensions indicates the importance of the 3-D flow simulations.

This work is currently in progress and the results will be presented in future reports and papers.

The completion of the hybrid model design tasks will lead to the construction of the three models 

representing the 20, 64 and 83 percent semispan stations of the full-scale wing. A series of ice-accretion 

test campaigns in the NASA IRT are planned in order to build the database of swept-wing ice accretion 

required for validation of icing simulation tools, fabrication of artificial ice shapes for aerodynamic 

testing and for CFD simulations of the iced swept wing.

2.4 Phase IV: High-Reynolds Number Aerodynamic Testing

High-Reynolds number aerodynamic testing will be conducted in Phase IV. A test campaign is 

planned for the ONERA F1 wind tunnel to investigate the aerodynamic effect of artificial ice shapes on 

the CRM65 wing. The F1 wind tunnel is a large-scale pressure tunnel having a test section 11.5 ft high by 

14.8 ft wide. A scaled, semispan model of the CRM65 wing will be designed and built to mount in the 

wind-tunnel force balance. A preliminary model sizing analysis has been carried out for the swept wing.

The main constraints on the geometry are the model span relative to the test-section height and the 

blockage defined as the ratio of the model planform area relative to the test-section area. While a large 

model is desirable to increase Reynolds number and the physical size of the artificial ice-shape 

geometries, a small model is desirable to mitigate tunnel-wall interference effects that can compromise 

the quality of the aerodynamic data. A sizing study was conducted to investigate these trade-offs. The best 

compromise was determined to be a 12.8 percent scale model with semispan that is 70 percent of the test-
NASA/TM—2013-216555 21

section height and planform area that is 8.5 percent of the test-section area. Dimensionally, this 

corresponds to a semispan of 8 ft and mean aerodynamic chord of 1.8 ft. The model aspect ratio and taper 

ratio are the same as for the full-scale CRM65 wing.

Aerodynamic testing will be performed up to a Reynolds number based upon mean aerodynamic chord 

of 10×106 and up to a Mach number of 0.3. In addition to standard force balance and surface pressure 

measurements, more detailed measurements are being investigated as described in Section 2.2.3. A series of 

full-span artificial ice shapes will be designed and built using the 3-D scan geometries from the IRT tests of 

the 20, 64 and 83 percent semispan stations of the CRM65 wing. One technical challenge associated with 

this effort is developing the full-span artificial ice shapes based upon the 3-D scan data from only three 

discrete spanwise stations tested in the IRT. It is anticipated that different lofting and interpolation methods 

will be used to design the artificial ice shapes. In addition, artificial ice shapes based upon the LEWICE3D 

simulation results may also be tested. A large number of configurations will be explored with varying levels 

of geometric fidelity. An important objective for the Phase IV research is determining the level of geometric 

fidelity required for the artificial ice shapes in order to yield accurate iced-wing aerodynamics. This will 

result in a large database of high-fidelity, iced-swept wing aerodynamic data at Reynolds numbers 

significantly higher than what is currently available in the public domain.

2.5 Phase V: Low-Reynolds Number Aerodynamic Testing

A series of low-Reynolds number aerodynamic test campaigns are planned for Phase V in order to 

develop a lower cost test capability for iced swept wings. There are several closed-return atmospheric wind 

tunnels in the United States with test section sizes of approximately 7 ft by 10 ft. One of these facilities will be 

selected for a series of low-Reynolds number aerodynamic test campaigns using a scaled, semispan model of 

the CRM65 wing that is very similar to that used for testing in the ONERA F1 wind tunnel. It is anticipated 

that the maximum Reynolds number based upon mean aerodynamic chord will be 2×106 at a Mach number of 

0.25. Aerodynamic measurements similar to Phase IV are planned. An important objective of Phase V is to 

quantify the differences in aerodynamic performance and key flowfield features between the low- and highReynolds number testing. If these differences prove to be small, the low-Reynolds number facility can be used 

for a variety of other investigations. For example, a series of ice-shape sensitivity tests could be conducted to 

explore the effects of ice feature size, shape and location making it possible to identify certain critical cases.

The low-Reynolds number facility could also be used for more detailed flowfield measurements using more 

sophisticated methods such as Particle Image Velocimetry (PIV).

2.6 Phase VI: High-Reynolds Number Validation Testing

For Phase VI, a second high-Reynolds number test campaign using the ONERA F1 wind tunnel is 

planned in order to conduct validation testing of critical ice shape cases identified in Phase V. In a large 

effort such as this, it is anticipated that all iced-wing configurations of interest cannot be tested in Phase 

IV. After conducting more detailed research at low-Reynolds number (in Phase V), it is expected that 

more ice-shape configurations of interest will be identified. It is also possible that significant Reynolds 

number effects will be identified. Therefore, further tests may be required at the end of this research effort 

in order to validate the final conclusions. The research conducted in this phase will “close-the-loop” on 

our understanding of swept-wing icing effects, generating an improved understanding and both 

experimental and computational simulation tools.

2.7 Phase VII: 3-D Ice Accretion and Flowfield Computational Simulation

In Phase VII, computational studies will be performed to identify the elements of the computational 

simulation process that are critical for accurate ice-shape predictions and reproduction of the 

aerodynamics of iced swept-wing geometries. This research is being conducted in parallel with the other 

tasks throughout the duration of the project. Computational research efforts will be undertaken using the 

geometric data of the CRM, ice shapes and the aerodynamic data from the experiments to evaluate current 

CFD capabilities with respect to simulation of iced swept-wing aerodynamics.
NASA/TM—2013-216555 22

Prior to any ice-accretion prediction, an accurate evaluation of the aerodynamic performance of the 

clean semispan wing as designed for the aerodynamic testing in Phases IV and V has to be performed, as it 

will be used as reference for the estimation of the performance degradation due to ice. Therefore, a first part 

of the computational activity will consider some parametric studies on the semispan clean wing in the windtunnel configuration. Among the different factors to be considered for simulation are clean aircraft flow 

characteristics, proper selection of grid resolution, turbulence models, and steady or unsteady calculations.

Other factors that can have a significant influence on the final solution (experimental set-up, Reynolds 

number effect, model deformations under loaded conditions due to pressurization) will also be investigated.

The added complexity of ice-shape geometries introduces the need to determine the degree of surface 

feature fidelity that can be and must be captured in the simulation to produce acceptable computational 

results. The evaluation of the aerodynamic performance of the wing with ice shapes will be mainly 

investigated using “standard” grid techniques (structured or unstructured meshes), with a detailed 

resolution up to the wall.

An initial look at what might be done with ice shape information from this program is show in Figure 

18. This figure demonstrates the process of using 3-D scan data from an actual ice shape as input to a 

standard grid generation software package. Figure 18(a) shows the full detail data from a scan of ice 

build-up on a NACA 23012 airfoil model. The discrete set of surface geometry coordinates from the laser 

scanner were post-processed to produce a watertight surface consisting of numerous triangles which were 

provided in a format compatible with most grid generation software. The surface description was then 

modified using smoothing algorithms within the post-processing software to obtain a lower resolution 

version of the surface, as seen in Figure 18(b). That file was then used as input to the grid generation 

software to produce the surface grid shown in Figure 18(c) and the volume grid in Figure 18(d). Different 

levels of smoothing will be examined to determine the influence on the aerodynamics very close to the 

surface as guidance for CFD modeling of the full CRM model geometry both with and without leadingedge ice shapes.

(a) (b)

(c) (d)

Figure 18.—Development of CFD grid from 3-D scanned ice shape.
NASA/TM—2013-216555 23

In parallel, ONERA will investigate the use of the “Immersed Boundary Conditions” (IBC) method in 

order to take the exact ice shape into account without having to mesh it explicitly (Ref. 59). A source term 

method will be introduced to accurately take into account the ice-shape boundaries. This IBC technique is 

currently under development in the ONERA elsA code (Ref. 60), and the validation phase for an 

application to iced wing shapes is planned to be carried out in the current project, according to the 

following sequence:

• Use of a 2-D configuration for preliminary evaluation, and development/improvement of tools for 

ice-shape problems; 

• Use a spanwise extension of the 2-D case to investigate the degree of geometrical refinement to 

be considered for ice shapes, and comparison with “standard” RANS computations in 2-D/2.5-D 

flow;

• Finally, 3-D applications on the configurations tested in F1, and comparison with “standard”

RANS computations in 3D flow.

Concerning ice-accretion prediction, ONERA and NASA will perform calculations using their own 

3-D icing suites. Initially, the ice shapes to be generated in the Phase III IRT test campaigns will be 

computed. The results of these simulations will be used for refining the definition of the test matrix and 

will be compared to the experimental results in a second step. ONERA will also work on the 

implementation of an automatic multistep procedure in its 3-D icing suite. With the current version of the 

suite (Ref. 61), the generation of new meshes for taking into account the ice-shape growth is done 

manually. This is complex and highly time consuming. Therefore the code is often used in a single step 

mode, which may not be accurate for complex ice shapes. The objective is to develop an automatic mesh 

adaptation and/or a re-meshing technique and to evaluate both its benefits and its range of application for 

the computation of 3-D ice shapes on a swept wing. A selected set of the tests performed in the IRT 

campaigns will be used for the assessment phase and comparisons with the NASA icing code will also be 

performed. 

3.0 Summary

Ice accretion and its aerodynamic effect on highly three-dimensional swept wings are extremely 

complex phenomena important to the continued design, certification and safe operation of small and large 

transport aircraft. Computational fluid dynamics codes have reached a level of maturity that they are 

being proposed by manufacturers for use in certification of aircraft for flight in icing. However, sufficient 

high-quality data to evaluate their performance on iced swept wings are not available. Thus, it is difficult 

to determine how much confidence can be placed in the results from CFD codes used in design, and 

particularly in certification. NASA, FAA, ONERA, the University of Illinois and Boeing, have embarked 

on a significant, collaborative research effort to address the technical challenges associated with sweptwing icing. The overall goal of ensuring continued flight safety will be achieved by improving the fidelity 

of experimental and computational simulation methods for swept-wing ice accretion formation and 

resulting aerodynamic effect. The purpose of this paper is to describe this research effort in more detail 

and report on the current results and status to date.

A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic 

experiments and computational simulation. Conducted in Phase I was a review of the technical literature 

associated with iced swept-wing aerodynamics with emphasis on classifying ice accretion based upon key 

aerodynamic features of the flowfield. This research has provided a framework for the icing and 

aerodynamic experiments to be conducted in later phases. The research conducted in Phase II has 

identified the 65 percent scale Common Research Model as the baseline, full-scale, swept-wing reference 

geometry utilized for the experiments and computational research. In addition, 3-D measurement 

techniques are being developed and validated to document, the experimental ice accretion geometries.
NASA/TM—2013-216555 24

Aerodynamic measurement techniques including pressure-sensitive paint and 5-hole probe wake surveys 

suitable to iced swept-wing experiments are being investigated. In Phase III, ice accretion testing will be 

conducted in the NASA IRT for three hybrid swept-wing models representing the 20, 64 and 83 percent

semispan stations of the baseline wing. Hybrid models preserve the full-scale, leading-edge geometry 

combined with a custom-designed truncated afterbody to reduce the overall model size and blockage in 

the IRT test section. A significant research effort is currently underway to develop the large-scale, hybrid 

swept-wing model design methods. Artificial ice shapes of varying geometric fidelity will be developed 

for aerodynamic testing over a large Reynolds number range in Phases IV and V. Exploring the effects of 

geometric fidelity and Reynolds number will require final validation tests in Phase VI. The ONERA F1 

pressurized wind-tunnel will be used for aerodynamic testing of a 12.8 percent scale semispan model of 

the baseline wing in Phases IV and VI. Data will be acquired for Reynolds numbers up to 10×106 and 

Mach numbers of 0.3. Lower-Reynolds number aerodynamic testing will also be conducted on a smaller 

scale semispan wing model in an atmospheric 7- by 10-ft size wind tunnel in Phase V. The research 

conducted in Phase VII will explore and further develop the use of computational simulation tools (CFD 

codes) for ice accretion and aerodynamics on swept wings.

The combined results of this research effort will result in an improved understanding of the icing on 

swept wings. The ice accretion testing will generate a large database of ice-accretion geometries on 

realistic large-scale swept wings. An additional outcome is the advanced method for designing hybrid 

icing wind-tunnel models. The aerodynamic testing will generate a large database of performance and 

flowfield information for swept wings with artificial ice shapes of varying geometric fidelity over a large 

range of Reynolds number. Concurrent computational research will be conducted to develop and validate 

ice accretion and aerodynamic simulation tools for large-scale swept wings. These databases will be 

available to manufacturers to evaluate the performance of CFD codes which they propose to use in the 

certification of aircraft for flight in icing conditions. These results will be used to determine the accuracy 

to which ice accretion must be predicted by computational tools or simulated in aerodynamic testing. For 

example, it is known for some cases that swept-wing ice accretion can be highly three dimensional (e.g., 

“scallops” or “lobster tails”). This research will help define how much detail of that three-dimensionality 

is critical to the aerodynamic performance and therefore must be accurately simulated.

References

1. Broeren, A.P., Addy, H.E., Jr., Bragg, M.B., Busch, G.T., Guffond, D., and Montreuil, E., 

“Aerodynamic Simulation of Ice Accretion on Airfoils,” NASA TP—2011-216929, June 2011.

2. Bragg, M.B., Broeren, A.P., Addy, H.E., Jr., Potapczuk, M., Guffond, D., and Montreuil, E., “Airfoil 

Ice-Accretion Aerodynamics Simulation,” AIAA Paper 2007-0085, Jan. 2007.

3. Bragg, M.B., Broeren, A.P., and Blumenthal, L.A., “Iced-Airfoil and Wing Aerodynamics,” SAE 

Paper 2003-01-2098, June 2003.

4. Bragg, M.B., Broeren, A.P., and Blumenthal, L.A., “Iced-Airfoil Aerodynamics,” Progress in 

Aerospace Sciences, Vol. 41, No. 5, July 2005, pp. 323-418.

5. Busch, G.T., Broeren, A.P., and Bragg, M.B., “Aerodynamic Simulation of a Horn-Ice Accretion on a 

Subscale Model,” AIAA Paper 2007-0087, Jan. 2007.

6. Busch, G.T., Broeren, A.P., and Bragg, M.B., “Aerodynamic Simulation of a Horn-Ice Accretion on a 

Subscale Model,” Journal of Aircraft, Vol. 45, No.2, Mar.-Apr. 2008, pp. 604-613.

7. Blumenthal, L.A., “Surface Pressure Measurements on a Three-Dimensional Ice Shape,” M.S. Thesis, 

Dept. of Aerospace Eng., Univ. of Illinois, Urbana, IL, 2005.

8. Busch, G.T., “Ice Accretion Aerodynamic Simulation on a Subscale Model, M.S. Thesis, Dept. of 

Aerospace Engineering, Univ. of Illinois, Urbana, IL, 2006.

9. Blumenthal, L.A., Busch, G.T., Broeren, A.P., and Bragg, M.B., “Issues in Ice Accretion 

Aerodynamic Simulation on a Subscale Model,” AIAA Paper 2006-0262, Jan. 2006.
NASA/TM—2013-216555 25

10. Busch, G.T., Broeren, A.P., and Bragg, M.B., “Aerodynamic Fidelity of Sub-scale Two-Dimensional 

Ice Accretion Simulation,” AIAA Paper 2008-7062, Aug. 2008.

11. Broeren, A.P., Busch, G.T., and Bragg, M.B., “Aerodynamic Fidelity of Ice Accretion Aerodynamic 

Simulation on a Subscale Model,” SAE 2007 Transactions: Journal of Aerospace, Vol. 116, Aug. 

2008, pp. 560-575; also SAE Paper 2007-01-3285, Sept. 2007.

12. Moens, F., “SUNSET Project: Numerical Investigations for the Preparation of the F1 Test 

Campaign,” ONERA Report No RT 1/12405 DAAP, April 2007.

13. CassouDeSalle, D. and Gilliot, A., “SUNSET (Studies on Scaling Effects due to Ice) Tests at High 

Reynolds Number in F1 Wind Tunnel,” ONERA Report No PV 4/12361 DSFM, December 2008.

14. CassouDeSalle, D., Gilliot, A., Geiler, C., Monnier, J.-C., Broeren, A.P., and Addy, H.E., Jr., 

“Experimental Investigations of Simulated Ice Accretions at High Reynolds Numbers in the Onera F1 

Wind Tunnel,” AIAA Paper 2009-4265, June 2009.

15. Broeren, A.P., Bragg, M.B., Addy, H.E., Jr., Lee, S., Moens, F., and Guffond, D., “Effect of HighFidelity Ice-Accretion Simulations on Full-Scale Airfoil Performance,” Journal of Aircraft, Vol. 47, 

No. 1, Jan.-Feb. 2010, pp. 240-254.

16. Broeren, A.P., Bragg, M.B., Addy, H.E., Jr., Lee, S., Moens, F., and Guffond, D., “Effect of HighFidelity Ice-Accretion Simulations on Full-Scale Airfoil Performance,” AIAA Paper 2008-0434, Jan. 

2008.

17. Busch, G.T., “Experimental Study of Full-Scale Iced-Airfoil Aerodynamic Performance using 

Subscale Simulations,” Ph.D. dissertation, Dept. of Aerospace Eng., Univ. of Illinois, Urbana, IL, 

2010.

18. Duclercq, M., Brunet, V., and Moens, F., “Physical Analysis of the Separated Flow Around an Iced 

Airfoil Based on ZDES Simulations,” AIAA Paper 2012-2798.

19. Vargas, M., “Current Experimental Basis for Modeling Ice Accretions on Swept Wings,” Journal of 

Aircraft, Vol. 44, No. 1. Jan-Feb 2007.

20. Vargas, M., “Swept Wing Icing Studies at NASA Glenn Research Center 1996-2006,” SAE Aircraft 

& Engine Icing International Conference, Seville, Spain, 2007, Paper No. 2007-01-3332.

21. Bidwell, C.S. and Mohler Jr., S.R., “Collection Efficiency and Ice Accretion Calculations for a 

Sphere, a Swept MS(1)-317 Wing, a Swept NACA-0012 Wing Tip, an Axisymmetric Inlet, and a 

Boeing 737-300 Inlet,” AIAA Paper 95-0755, Jan. 1995.

22. Bidwell, C.S. and Potapczuk, M.G., “User’s Manual for the NASA Lewis Three-Dimensional Ice 

Accretion Code (LEWICE3D),” NASA TM–105974, Dec. 1993.

23. Potapczuk, M.G. and Bidwell, C.S., “Swept Wing Ice Accretion Modeling,” NASA TM–103114, Jan. 

1990.

24. Potapczuk, M.G. and Bidwell, C.S., “Numerical Simulation of Ice Growth on a MS-371 Swept Wing 

Geometry,” NASA TM 103705, Jan. 1991.

25. Potapczuk, M., Papadakis, M. and Vargas, M., “LEWICE Modeling of Swept Wing Ice Accretions,”

AIAA Paper 2003-6565, Jan. 2003.

26. Hedde, T., and Guffond, D., “Development of a Three-Dimensional Icing Code, Comparison with 

Experimental Shapes,” AIAA Paper 92-0041, Jan. 1992.

27. Hedde, T., and Guffond, D., “Improvement of the ONERA 3D Icing Code Comparison with 3D 

Experimental Shapes,” AIAA Paper 93-0169, Jan. 1993.

28. Szilder, K., McIlwain, S., and Lozowski, E.P., “Numerical Simulation of Complex Ice Shapes on 

Swept Wings,” ICAS Paper 2006-2.5.1, Sept. 2006.

29. Presteau, X., Montreuil, E., Leroy, A., Guffond, D., Henry, R., and Personne, P, “Experimental Study 

of the Scallop Formation on Swept Cylinder,” SAE Paper 2007-01-3296, Sept. 2007.

30. Presteau, X., Montreuil, E., Chazottes, A., Vancassel, X., and Personne, P, “Experimental and 

Numerical Study of Scallop Ice on Swept Cylinder,” AIAA Paper 2009-4124, June 2009.

31. Bragg, M.B., Kerho, M.F., and Khodadoust, A., “LDV Flowfield Measurements on a Straight and 

Swept Wing with a Simulated Ice Accretion,” AIAA Paper 93-0300, Jan. 1993.
NASA/TM—2013-216555 26

32. Kwon, O.J., and Sankar, L.N., “Numerical Study of the Effects of Icing on Fixed and Rotary Wing 

Performance,” AIAA Paper 91-0662, Jan. 1990.

33. Khodadoust, A., and Bragg, M.B, “Measured Aerodynamic Performance of a Swept Wing with a

Simulated Ice Accretion,” AIAA Paper 90-0490, Jan. 1990.

34. Bragg, M.B., Khodadoust, A., Soltani, R., Wells, S., Kerho, M., “Effect of a Simulated Ice Accretion 

on the Aerodynamics of a Swept Wing,” AIAA Paper 91-0442, Jan. 1991.

35. Vargas, M., Papadakis, M., Potapczuk, M., Addy, H., Sheldon, D. and Giriunas, J., “Ice Accretions 

on a Swept GLC-305 Airfoil,” SAE Paper 02GAA-43 and NASA TM-2002-211557, April 2002.

36. Papadakis, M., Yeong, H. W., Won, S. C., Vargas, M. and Potapczuk, M., “Aerodynamic 

Performance of a Swept Wing with Ice Accretions,” AIAA Paper 2003-0731, Jan. 2003.

37. Papadakis, M., Yeong, H.W., Wong, S, Vargas, M and Potapczuk, M., “Experimental Investigation of 

Ice Accretion Effects on a Swept Wing,” DOT/FAA/AR-05/39, Aug. 2005.

38. Reehorst, A.L., Potapczuk, M.G., Ratvasky, T.P., and Gile-Laflin, B.E., “Wind Tunnel Measured 

Effects on a Twin-Engine Short-Haul Transport Caused by Simulated Ice Accretions,” AIAA Paper 

96-0871, also NASA TM 107143, Jan. 1996.

39. Reehorst, A.L., Potapczuk, M.G., Ratvasky, T.P., and Gile-Laflin, B.E., “Wind Tunnel Measured 

Effects on a Twin-Engine Short-Haul Transport Caused by Simulated Ice Accretions Data Report,” 

NASA TM–107419, May 1997.

40. Papadakis, M., Yeong, H.W., Chandrasekharan, R., Hinson, M., Ratvasky, T., and Giriunas, J., 

“Experimental Investigation of Simulated Ice Accretions on a Full-Scale T-tail,” AIAA Paper 2001-

0090, Jan. 2001.

41. Papadakis, M., Yeong, H.W., Chandrasekharan, R., Hinson, M., and Ratvasky, T. “Effects of 

Roughness on the Aerodynamic Performance of a Business Jet Tail,” AIAA Paper 2002-0242, Jan. 

2002.

42. Broeren, A.P., Diebold, J.M., and Bragg, M.B., “Aerodynamic Classification of Swept-Wing Ice 

Accretion,” NASA/TM—2013-216381, May 2013, also DOT/FAA/TC-13/21.

43. Diebold, J.M., Broeren, A.P. and Bragg, M.B., “Aerodynamic Classification of Swept-Wing Ice 

Accretion,” AIAA 5th Atmospheric and Space Environments Conference, San Diego, CA, June 24-

27, 2013 (submitted for publication).

44. Vassberg, J.C., DeHann, M.A., Rivers, S.M., and Wahls, R.A., “Development of a Common 

Research Model for Applied CFD Validation Studies,” AIAA Paper 2008-6919, Aug. 2008.

45. Rivers, M.B., and Dittberner, A., “Experimental Investigation of the NASA Common Research 

Model,” AIAA Paper 2010-4218, June 2010.

46. Rivers, M.B., and Dittberner, A., “Experimental Investigations of the NASA Common Research 

Model in the NASA Langley National Transonic Facility and the NASA Ames 11-Ft Transonic Wind 

Tunnel,” AIAA Paper 2011-1126, Jan. 2011. 

47. Vassberg, J. C., Tinoco, E. N., Mani, M., Rider, B., Zickuhr, T., Levy, D.W., Brodersen, O., Eisfeld, 

B., Crippa, S., Wahls, R. A., Morrison, J. H., Mavriplis, D.J., and Murayama, M., “Summary of the 

Fourth AIAA CFD Drag Prediction Workshop,” AIAA Paper 2010-4547, June 2010.

48. Bosetti, C., Paul, B.P., Jr., and Malone, A.M., “Ice Shape Characterization to Aid in Replicating Ice 

Shapes for Subsequent Analysis,” AIAA Paper 2010-7534, Aug. 2010.

49. Reehorst, A.L., and Richter, G.P., “New Methods and Materials for Molding and Casting Ice 

Formations,” NASA TM-100126, Sept. 1987. 

50. Lee, S., Broeren, A.P., Addy, H.E., Jr., Sills, R., and Pifer, E.M., “Development of 3-D Ice Accretion 

Measurement Method,” AIAA Paper 2012-2938, June 2012, also NASA/TM—2012-217702, Sept. 

2012.

51. Diebold, J.M., Monastero, M.C., and Bragg, M.B., “Aerodynamics of a Swept-Wing with Ice 

Accretion at Low Reynolds Number,” AIAA Paper 2012-2795, June 2012.

52. Diebold, J.M., “Aerodynamics of a Swept Wing with Leading-Edge Ice at Low Reynolds Number”

M.S. Thesis, Department of Aerospace Engineering, University of Illinois, Urbana, IL, Aug. 2012.
NASA/TM—2013-216555 27

53. Diebold, J.M., and Bragg, M.B., “Study of a Swept-Wing with Leading-Edge Ice Using a Wake 

Survey Technique,” AIAA Paper 2013-0245, Jan. 2013.

54. Sclafani, A.J., Slotnick, J.P., Vassberg, J.C., and Pulliam, T.H., “Extended OVERFLOW Analysis of 

the NASA Trap Wing Wind Tunnel Model,” AIAA Paper 2012-2919, June 2012.

55. Saeed, F., Selig, M.S., and Bragg, M.B., “Hybrid Airfoil Design Procedure Validation for Full-Scale 

Ice Accretion Simulation,” Journal of Aircraft, Vol. 36, No. 5, Sept.-Oct., 1999, pp. 769-776.

56. Saeed, F., Selig, M.S., and Bragg, M.B., “Hybrid Airfoil Design to Simulate Full-Scale Ice Accretion 

Throughout a Given α Range,” Journal of Aircraft, Vol. 35, No. 2, Mar.-Apr., 1998, pp. 233-239.

57. Saeed, F., Selig, M.S., and Bragg, M.B., “Design of Subscale Airfoils with Full-Scale Leading Edges 

for Ice Accretion Testing,” Journal of Aircraft, Vol. 34, No. 1, Jan.-Feb., 1997, pp. 94-100.

58. Fujiwara, G.E.C., Woodard, B.S., Wiberg, B.D., Mortonson, A.J., Bragg, M.B., “A Hybrid Airfoil 

Design Method for Icing Wind Tunnel Tests,” AIAA 5th Atmospheric and Space Environments 

Conference, San Diego, CA, June 24-27, 2013 (submitted for publication).

59. Terracol M., Manoha E., and Lemoine B., “Investigation of the Unsteady Flow and Noise Sources 

Generation in a Slat Cove: Hybrid Zonal RANS/LES Simulation and Dedicated Experiment,” AIAA 

Paper 2011-3203, 20th AIAA Computational Fluid Dynamics Conference 2011, Honolulu, HI, 27-30, 

June 2011.

60. Reneaux J., Beaumier P. and Giroudroux-Lavigne P. “Advanced Aerodynamic Applications with the 

elsA Software,” The Onera Journal, AerospaceLab Journal, Issue 2, March 2011. 

(www.aerospacelab-journal.org).

61. Villedieu, P., Bobo, D., Guffond, D., Trontin, P. “SLD Lagrangian Modelling and Capability 

Assessment in the Frame of ONERA 3D Icing Suite,” AIAA Paper 2012-3132, New Orleans, 2012.